1. Field of the Invention
The present invention relates generally to spacecraft and, more particularly, to acquiring and determining spacecraft attitudes.
2. Description of the Related Art
FIG. 1 is a diagram 20 that illustrates exemplary orbits of a spacecraft about an Earth 22. The spacecraft is shown both as a spacecraft 24D with its solar wings 25 in a deployed configuration wherein they extend from its body 26 and as a spacecraft 24S with its solar wings in a stowed configuration wherein they adjoin the body. The spacecraft is initially launched along a launch path 27 into a parking orbit 28 prior to its installation into a final orbit such as a geosynchronous orbit (GEO) 30. Transfer between the parking orbit 28 and the GEO 30 is realized along transfer orbits such as exemplary transfer orbits 32 and 34.
The elliptical transfer orbit 32 has a perigee 36 tangent to the parking orbit 28 and an apogee 37 tangent to the GEO 30. Insertion of the spacecraft into and out of the transfer orbit 32 is typically accomplished with a motor that is generally referred to as an apogee motor which is of a motor type (e.g., solid propellant motor and bi-propellant liquid motor) that can realize a large thrust for a short time. The apogee motor is fired at the perigee 36 to attain a spacecraft velocity appropriate at that altitude for the transfer orbit 32 and is fired again at the apogee 37 to attain a spacecraft velocity appropriate at that altitude for the GEO 30.
Because the apogee motor does not require large amounts of electrical power, the solar wings 25 are typically in a stowed configuration during the transfer orbit 32 and are then extended into a deployed configuration in the GEO 30. The stowed configuration also protects the fragile solar wings from the acceleration of the apogee motor and from contamination by propellants of the apogee motor.
In contrast to the transfer orbit 32, the transfer orbit 34 has an apogee 38 whose altitude exceeds that of the GEO 30. This transfer orbit is typically converted to the GEO 30 by continuous firing of a low thrust engine such as an ion propulsion thruster. A thruster of this type must generate large electrostatic fields over long periods of time and, accordingly, the solar wings are in their deployed configuration during the transfer orbit 34 to provide the necessary electrical power. They are not in danger of being damaged during this transfer because the ion thrust is extremely low.
The attitude of the spacecraft 24D is typically defined with reference to a body-fixed system of three orthogonal axes which are shown in the enlarged view of FIG. 2A to be a yaw axis 42, a roll axis 43 and a pitch axis 44. The axes are fixed relative to the spacecraft""s body 26 and their arrowheads indicate the generally-accepted positive directions of the axes. The solar wings 25 extend from opposite sides of the body 26, rotate about the pitch axis 43 and carry an array 40 of solar cells on one surface to facilitate power generation. Because the solar wings generally have considerable length, they are illustrated in a shortened form in FIG. 2A.
When the spacecraft is in the GEO 30 of FIG. 1, its attitude must be carefully controlled to maintain it in a xe2x80x9cservice attitudexe2x80x9d that permits it to carry out the service operations for which it was designed. An exemplary service attitude directs the yaw axis 42 at the Earth 22 with the roll axis 43 in the plane of the GEO and the pitch axis 44 orthogonal to the GEO plane.
Various anomalies in a spacecraft""s attitude control system can cause it to depart from its service attitude and, further, to take on an unknown attitude in which its attitude sensors (e.g., star trackers) fail to provide attitude information. To prevent failure of the spacecraft and its operations, the spacecraft must promptly acquire an attitude in which its solar wings generate sufficient power to maintain a viable spacecraft. In addition, the spacecraft must determine its attitude so that it can be subsequently urged to its service attitude.
Spacecraft and their operation are generally expensive endeavors so that loss of the service attitude is of great concern. In a communication spacecraft, for example, revenues and customers are lost when the spacecraft""s service is interrupted. The spacecraft""s xe2x80x9creturn-to-servicexe2x80x9d time must be reduced to minimize these costs. When the service attitude is lost, it is therefore important to not only acquire and determine a power-safe attitude but to do it promptly.
Although the term xe2x80x9cservicexe2x80x9d is typically applied to operations conducted in a spacecraft""s permanent orbit, it is used herein to also indicate service operations during a preliminary orbit such as the transfer orbit 32 of FIG. 1. When it is in this orbit, the enlarged view of FIG. 2B shows that the solar wings of the spacecraft 24S are stowed to adjoin opposite sides of its body 25 and arranged so that a portion 46, of each solar cell array is parallel to the yaw axis 42. Because the spacecraft can be permanently lost if its attitude is not properly controlled throughout the transfer orbit 32, it is important to promptly acquire a power-safe attitude when in the transfer orbit 32.
Various methods have been proposed for acquiring and determining spacecraft attitudes. A method for determining the instantaneous attitude of a spinning spacecraft, for example, is disclosed in U.S. Pat. No. 5,020,744 (issued on Jun. 4, 1991 to Schwarzschild). The method requires inputs from a sun sensor, an earth sensor and a 3-axis gyroscope assembly. U.S. Pat. No. 5,255,879 (issued on Oct. 26, 1993 to Yocum, et al.) provides a method for directing the roll axis of a spacecraft along the sun line but it requires that a spacecraft carry three single-axis sun sensors.
A method for determining a spacecraft""s attitude is provided by U.S. Pat. No. 5,412,574 (issued on May 2, 1995 to Bender, et al.). This method requires a terrestrial sensor (e.g., an earth sensor or a beacon sensor) and at least one star tracker or a cross-link sensor. U.S. Pat. No. 5,597,142 (issued Jan. 28, 1997 to Leung, et al.) teaches the use of a sun sensor, an earth sensor and a 3-axis gyroscope assembly to obtain a desired spacecraft attitude. Finally, U.S. Pat. No. 5,865,402 (issued Feb. 2, 1999 to Fischer et al.) discloses a method of acquiring a spacecraft attitude with a sun sensor, an earth sensor and a direction vector measurement device such as a star sensor or a magnetometer.
As evidenced in these examples, conventional methods for acquiring and determining spacecraft attitudes have typically:
a) required numerous attitude sensors which must often be added to those used in other spacecraft operations,
b) required numerous sequential maneuvers, and
c) reached a power, safe attitude that significantly departs from the spacecraft""s service attitude and, therefore, lengthens the return-to-service time.
Because weight and space are limited assets in spacecraft and because increased return-to-service time causes loss of revenue, it is of significant importance to provide improved methods for acquiring and determining power-safe spacecraft attitudes.
The present invention is of particular use in spacecraft that have, for any reason, lost the spacecraft""s service attitude that permits it to carry out the service operations for which it was designed. As part of this loss of service attitude, the spacecraft has typically also lost knowledge of its attitude, i.e., it cannot determine its attitude. In these cases, it is imperative that the spacecraft is quickly returned to its service attitude to minimize loss of revenue and, in extreme cases, loss of the spacecraft.
The present invention provides methods and structures for acquiring and determining a xe2x80x9cpower-safe attitudexe2x80x9dxe2x80x94that being an attitude in which wing current is sufficient to support the spacecraft""s housekeeping operations and from which, the spacecraft can be subsequently returned to its service attitude.
A method embodiment includes the steps of
a) at a rotation rate, rotating a solar wing about a wing axis;
b) at a slew rate that is slower than the rotation rate, slewing the spacecraft about a slew axis that is tilted at least 30 degrees from the wing axis; and
c) continuing at least one of the rotating and slewing steps to acquire a power-safe attitude in which the wing current exceeds a wing-current threshold and a star tracker""s field-of-view contains an identifiable set of stars which determines the power-safe attitude.
Methods of the invention are particularly useful because they:
a) require only a single star tracker which is an attitude sensor that is often included in a spacecraft for other purposes in such cases, the invention does not require the addition of any sensors,
b) monitor a simple parameter (i.e., wing current) to determine solar-wing power generation,
c) comprise simple maneuvers, and
d) in many cases, acquire a power-safe attitude that does not significantly differ from the spacecraft""s service attitude to thereby reduce the spacecraft""s return-to-service time.